Turbine section of gas turbine engine

ABSTRACT

A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor including a circumferential array of blades, a low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area and a low pressure turbine section. The low pressure turbine section includes a maximum gas path radius, the blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the blades is equal to or greater than 0.35, and is less than 0.55.

CROSS-REFERENCE TO RELATED APPLICATION

The present application is a continuation of U.S. patent applicationSer. No. 17/062756 filed Oct. 5, 2020, which is a continuation of U.S.patent application Ser. No. 16/025094 filed Jul. 2, 2018, which is acontinuation of U.S. patent application Ser. No. 15/292,472, filed Oct.13, 2016, which is a continuation of U.S. patent application Ser. No.14/793,785, filed Jul. 8, 2015, which is a continuation-in-part of U.S.patent application Ser. No. 14/692,090, filed Apr. 21, 2015, which is acontinuation of U.S. patent application Ser. No. 13/599,175, filed Aug.30, 2012, which is a continuation of U.S. patent application Ser. No.13/475,252, now U.S. Pat. No. 8,844,265, issued Sep. 30, 2014, filed May18, 2012. U.S. patent application Ser. No. 13/475,252 is acontinuation-in-part of U.S. patent application Ser. No. 11/832,107,filed Aug. 1, 2007, and claims the benefit of U.S. Patent ProvisionalApplication No. 61/593,190, filed Jan. 31, 2012, and U.S. ProvisionalApplication No. 61/498,516, filed Jun. 17, 2011.

BACKGROUND

The disclosure relates to turbofan engines. More particularly, thedisclosure relates to low pressure turbine sections of turbofan engineswhich power the fans via a speed reduction mechanism.

There has been a trend toward increasing bypass ratio in gas turbineengines. This is discussed further below. There has generally been acorrelation between certain characteristics of bypass and the diameterof the low pressure turbine section sections of turbofan engines.

SUMMARY

One aspect of the disclosure involves a turbofan engine having an enginecase and a gaspath through the engine case. A fan has a circumferentialarray of fan blades. The engine further has a compressor in fluidcommunication with the fan, a combustor in fluid communication with thecompressor, a turbine in fluid communication with the combustor, whereinthe turbine includes a low pressure turbine section having 3 to 6 bladestages. A speed reduction mechanism couples the low pressure turbinesection to the fan. A bypass area ratio is greater than about 6.0. Aratio of the total number of airfoils in the low pressure turbinesection divided by the bypass area ratio is less than about 170, saidlow pressure turbine section airfoil count being the total number ofblade airfoils and vane airfoils of the low pressure turbine section.

In additional or alternative embodiments of any of the foregoingembodiments, the bypass area ratio may be greater than about 8.0 or maybe between about 8.0 and about 20.0.

In additional or alternative embodiments of any of the foregoingembodiments, a fan case may encircle the fan blades radially outboard ofthe engine case.

In additional or alternative embodiments of any of the foregoingembodiments, the compressor may comprise a low pressure compressorsection and a high pressure compressor section.

In additional or alternative embodiments of any of the foregoingembodiments, the blades of the low pressure compressor section and lowpressure turbine section may share a low shaft.

In additional or alternative embodiments of any of the foregoingembodiments, the high pressure compressor section and a high pressureturbine section of the turbine may share a high shaft.

In additional or alternative embodiments of any of the foregoingembodiments, there are no additional compressor or turbine sections.

In additional or alternative embodiments of any of the foregoingembodiments, the speed reduction mechanism may comprise an epicyclictransmission coupling the low speed shaft to a fan shaft to drive thefan with a speed reduction.

In additional or alternative embodiments of any of the foregoingembodiments, the low pressure turbine section may have an exemplary 2 to6 blade stages or 2 to 3 blade stages.

In additional or alternative embodiments of any of the foregoingembodiments, a hub-to-tip ratio (R_(I):R_(O)) of the low pressureturbine section may be between about 0.4 and about 0.5 measured at themaximum R_(O) axial location in the low pressure turbine section.

In additional or alternative embodiments of any of the foregoingembodiments, a ratio of maximum gaspath radius along the low pressureturbine section to maximum radius of the fan may be less than about0.55, or less than about 0.50, or between about 0.35 and about 0.50.

In additional or alternative embodiments of any of the foregoingembodiments, the ratio of low pressure turbine section airfoil count tobypass area ratio may be between about 10 and about 150.

In additional or alternative embodiments of any of the foregoingembodiments, the airfoil count of the low pressure turbine section maybe below about 1600.

In additional or alternative embodiments of any of the foregoingembodiments, the engine may be in combination with a mountingarrangement (e.g., of an engine pylon) wherein an aft mount reacts atleast a thrust load.

The details of one or more embodiments are set forth in the accompanyingdrawings and the description below. Other features, objects, andadvantages will be apparent from the description and drawings, and fromthe claims.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is an axial sectional view of a turbofan engine.

FIG. 2 is an axial sectional view of a low pressure turbine section ofthe engine of FIG. 1.

FIG. 3 is transverse sectional view of transmission of the engine ofFIG. 1.

FIG. 4 shows another embodiment.

FIG. 5 shows yet another embodiment.

Like reference numbers and designations in the various drawings indicatelike elements.

DETAILED DESCRIPTION

FIG. 1 shows a turbofan engine 20 having a main housing (engine case) 22containing a rotor shaft assembly 23. An exemplary engine is ahigh-bypass turbofan. In such an engine, the normal cruise conditionbypass area ratio of air mass flowing outside the case 22 (e.g., thecompressor sections and combustor) to air mass passing through the case22 is typically in excess of about 4.0 and, more narrowly, typicallybetween about 4.0 and about 12.0. Via high 24 and low 25 shaft portionsof the shaft assembly 23, a high pressure turbine section (gasgenerating turbine) 26 and a low pressure turbine section 27respectively drive a high pressure compressor section 28 and a lowpressure compressor section 30. As used herein, the high pressureturbine section experiences higher pressures that the low pressureturbine section. A low pressure turbine section is a section that powersa fan 42. Although a two-spool (plus fan) engine is shown, one of manyalternative variations involves a three-spool (plus fan) engine whereinan intermediate spool comprises an intermediate pressure compressorbetween the low fan and high pressure compressor section and anintermediate pressure turbine between the high pressure turbine sectionand low pressure turbine section.

The engine extends along a longitudinal axis 500 from a fore end to anaft end. Adjacent the fore end, a shroud (fan case) 40 encircles the fan42 and is supported by vanes 44. An aerodynamic nacelle around the fancase is shown and an aerodynamic nacelle 45 around the engine case isshown.

The low shaft portion 25 of the rotor shaft assembly 23 drives the fan42 through a speed reduction mechanism 46. An exemplary speed reductionmechanism is an epicyclic transmission, namely a star or planetary gearsystem. As is discussed further below, an inlet airflow 520 entering thenacelle is divided into a portion 522 passing along a core flowpath 524and a bypass portion 526 passing along a bypass flowpath 528. With theexception of diversions such as cooling air, etc., flow along the coreflowpath sequentially passes through the low pressure compressorsection, high pressure compressor section, a combustor 48, the highpressure turbine section, and the low pressure turbine section beforeexiting from an outlet 530.

FIG. 3 schematically shows details of the transmission 46. A forward endof the low shaft 25 is coupled to a sun gear 52 (or other high speedinput to the speed reduction mechanism). The externally-toothed sun gear52 is encircled by a number of externally-toothed star gears 56 and aninternally-toothed ring gear 54. The exemplary ring gear is coupled tothe fan to rotate with the fan as a unit.

The star gears 56 are positioned between and enmeshed with the sun gearand ring gear. A cage or star carrier assembly 60 carries the star gearsvia associated journals 62. The exemplary star carrier is substantiallyirrotatably mounted relative via fingers 404 to the case 22.

Another transmission/gearbox combination has the star carrier connectedto the fan and the ring is fixed to the fixed structure (case) ispossible and such is commonly referred to as a planetary gearbox.

The speed reduction ratio is determined by the ratio of diameters withinthe gearbox. An exemplary reduction is between about 2:1 and about 13:1.

The exemplary fan (FIG. 1) comprises a circumferential array of blades70. Each blade comprises an airfoil 72 having a leading edge 74 and atrailing edge 76 and extending from an inboard end 78 at a platform toan outboard end 80 (i.e., a free tip). The outboard end 80 is in closefacing proximity to a rub strip 82 along an interior surface 84 of thenacelle and fan case.

To mount the engine to the aircraft wing 92, a pylon 94 is mounted tothe fan case and/or to the other engine cases. The exemplary pylon 94may be as disclosed in U.S. patent application Ser. No. 11/832,107(US2009/0056343A1). The pylon comprises a forward mount 100 and anaft/rear mount 102. The forward mount may engage the engine intermediatecase (IMC) and the aft mount may engage the engine thrust case. The aftmount reacts at least a thrust load of the engine.

To reduce aircraft fuel burn with turbofans, it is desirable to producea low pressure turbine with the highest efficiency and lowest weightpossible. Further, there are considerations of small size (especiallyradial size) that benefit the aerodynamic shape of the engine cowlingand allow room for packaging engine subsystems.

FIG. 2 shows the low pressure turbine section 27 as comprising anexemplary three blade stages 200, 202, 204. An exemplary blade stagecount is 2-6, more narrowly, 2-4, or 2-3, 3-5, or 3-4. Interspersedbetween the blade stages are vane stages 206 and 208. Each exemplaryblade stage comprises a disk 210, 212, and 214, respectively. Acircumferential array of blades extends from peripheries of each of thedisks. Each exemplary blade comprises an airfoil 220 extending from aninner diameter (ID) platform 222 to an outer diameter (OD) shroud 224(shown integral with the airfoil

An alternative may be an unshrouded blade with a rotational gap betweenthe tip of the blade and a stationary blade outer air seal (BOAS). Eachexemplary shroud 224 has outboard sealing ridges which seal withabradable seals (e.g., honeycomb) fixed to the case. The exemplary vanesin stages 206 and 208 include airfoils 230 extending from ID platforms232 to OD shrouds 234. The exemplary OD shrouds 234 are directly mountedto the case. The exemplary platforms 232 carry seals for sealing withinter-disk knife edges protruding outwardly from inter-disk spacerswhich may be separate from the adjacent disks or unitarily formed withone of the adjacent disks.

Each exemplary disk 210, 212, 214 comprises an enlarged central annularprotuberance or “bore” 240, 242, 244 and a thinner radial web 246, 248,250 extending radially outboard from the bore. The bore impartsstructural strength allowing the disk to withstand centrifugal loadingwhich the disk would otherwise be unable to withstand.

A turbofan engine is characterized by its bypass ratio (mass flow ratioof air bypassing the core to air passing through the core) and thegeometric bypass area ratio (ratio of fan duct annulus areaoutside/outboard of the low pressure compressor section inlet (i.e., atlocation 260 in FIG. 1) to low pressure compressor section inlet annulusarea (i.e., at location 262 in FIG. 2). High bypass engines typicallyhave bypass area ratio of at least four. There has been a correlationbetween increased bypass area ratio and increased low pressure turbinesection radius and low pressure turbine section airfoil count. As isdiscussed below, this correlation may be broken by having an engine withrelatively high bypass area ratio and relatively low turbine size.

By employing a speed reduction mechanism (e.g., a transmission) to allowthe low pressure turbine section to turn very fast relative to the fanand by employing low pressure turbine section design features for highspeed, it is possible to create a compact turbine module (e.g., whileproducing the same amount of thrust and increasing bypass area ratio).The exemplary transmission is a epicyclic transmission. Alternativetransmissions include composite belt transmissions, metal chain belttransmissions, fluidic transmissions, and electric means (e.g., amotor/generator set where the turbine turns a generator providingelectricity to an electric motor which drives the fan).

Compactness of the turbine is characterized in several ways. Along thecompressor and turbine sections, the core gaspath extends from aninboard boundary (e.g., at blade hubs or outboard surfaces of platformsof associated blades and vanes) to an outboard boundary (e.g., at bladetips and inboard surfaces of blade outer air seals for unshrouded bladetips and at inboard surfaces of OD shrouds of shrouded blade tips and atinboard surfaces of OD shrouds of the vanes). These boundaries may becharacterized by radii R_(I) and R_(O), respectively, which vary alongthe length of the engine.

For low pressure turbine radial compactness, there may be a relativelyhigh ratio of radial span (R_(O)-R_(I)) to radius (R_(O) or R_(I)).Radial compactness may also be expressed in the hub-to-tip ratio(R_(I):R_(O)). These may be measured at the maximum R_(O) location inthe low pressure turbine section. The exemplary compact low pressureturbine section has a hub-to-tip ratio close to about 0.5 (e.g., about0.4-0.5 or about 0.42-0.48, with an exemplary about 0.46).

Another characteristic of low pressure turbine radial compactness isrelative to the fan size. An exemplary fan size measurement is themaximum tip radius R_(Tmax) of the fan blades. An exemplary ratio is themaximum R_(O) along the low pressure turbine section to R_(Tmax) of thefan blades. Exemplary values for this ratio are less than about 0.55(e.g., about 0.35-55), more narrowly, less than about 0.50, or about0.35-0.50.

To achieve compactness the designer may balance multiple physicalphenomena to arrive at a system solution as defined by the low pressureturbine hub-to-tip ratio, the fan maximum tip radius to low pressureturbine maximum R_(O) ratio, the bypass area ratio, and the bypass arearatio to low pressure turbine airfoil count ratio. These concernsinclude, but are not limited to: a) aerodynamics within the low pressureturbine, b) low pressure turbine blade structural design, c) lowpressure turbine disk structural design, and d) the shaft connecting thelow pressure turbine to the low pressure compressor and speed reductiondevice between the low pressure compressor and fan. These physicalphenomena may be balanced in order to achieve desirable performance,weight, and cost characteristics.

The addition of a speed reduction device between the fan and the lowpressure compressor creates a larger design space because the speed ofthe low pressure turbine is decoupled from the fan. This design spaceprovides great design variables and new constraints that limitfeasibility of a design with respect to physical phenomena. For examplethe designer can independently change the speed and flow area of the lowpressure turbine to achieve optimal aerodynamic parameters defined byflow coefficient (axial flow velocity/wheel speed) and work coefficient(wheel speed/square root of work). However, this introduces structuralconstraints with respect blade stresses, disk size, material selection,etc.

In some examples, the designer can choose to make low pressure turbinesection disk bores much thicker relative to prior art turbine bores andthe bores may be at a much smaller radius RB. This increases the amountof mass at less than a “self sustaining radius”. Another means is tochoose disk materials of greater strength than prior art such as the useof wrought powdered metal disks to allow for extremely high centrifugalblade pulls associated with the compactness.

Another variable in achieving compactness is to increase the structuralparameter AN² which is the annulus area of the exit of the low pressureturbine divided by the low pressure turbine rpm squared at its redlineor maximum speed. Relative to prior art turbines, which are greatlyconstrained by fan blade tip mach number, a very wide range of AN²values can be selected and optimized while accommodating suchconstraints as cost or a countering, unfavorable trend in low pressureturbine section shaft dynamics. In selecting the turbine speed (andthereby selecting the transmission speed ratio, one has to be mindfulthat at too high a gear ratio the low pressure turbine section shaft(low shaft) will become dynamically unstable.

The higher the design speed, the higher the gear ratio will be and themore massive the disks will become and the stronger the low pressureturbine section disk and blade material will have to be. All of theseparameters can be varied simultaneously to change the weight of theturbine, its efficiency, its manufacturing cost, the degree ofdifficulty in packaging the low pressure turbine section in the corecowling and its durability. This is distinguished from a prior artdirect drive configuration, where the high bypass area ratio can only beachieved by a large low pressure turbine section radius. Because thatradius is so very large and, although the same variables (airfoilturning, disk size, blade materials, disk shape and materials, etc.) aretheoretically available, as a practical matter economics and engine fuelburn considerations severely limit the designer's choice in theseparameters.

Another characteristic of low pressure turbine section size is airfoilcount (numerical count of all of the blades and vanes in the lowpressure turbine). Airfoil metal angles can be selected such thatairfoil count is low or extremely low relative to a direct driveturbine. In known prior art engines having bypass area ratio above 6.0(e.g. 8.0-20), low pressure turbine sections involve ratios of airfoilcount to bypass area ratio above 190.

With the full range of selection of parameters discussed aboveincluding, disk bore thickness, disk material, hub to tip ratio, andR_(O)/R_(Tmax), the ratio of airfoil count to bypass area ratio may bebelow about 170 to as low as 10. (e.g., below about 150 or an exemplaryabout 10-170, more narrowly about 10-150). Further, in such embodimentsthe airfoil count may be below about 1700, or below about 1600.

FIG. 4 shows an embodiment 600, wherein there is a fan drive turbine 608driving a shaft 606 to in turn drive a fan rotor 602. A gear reduction604 may be positioned between the fan drive turbine 608 and the fanrotor 602. This gear reduction 604 may be structured and operate likethe gear reduction disclosed above. A compressor rotor 610 is driven byan intermediate pressure turbine 612, and a second stage compressorrotor 614 is driven by a turbine rotor 216. A combustion section 618 ispositioned intermediate the compressor rotor 614 and the turbine section616.

FIG. 5 shows yet another embodiment 700 wherein a fan rotor 702 and afirst stage compressor 704 rotate at a common speed. The gear reduction706 (which may be structured as disclosed above) is intermediate thecompressor rotor 704 and a shaft 708 which is driven by a low pressureturbine section.

One or more embodiments have been described. Nevertheless, it will beunderstood that various modifications may be made. For example, whenreengineering from a baseline engine configuration, details of thebaseline may influence details of any particular implementation.Accordingly, other embodiments are within the scope of the followingclaims.

What is claimed is:
 1. A gas turbine engine comprising: a propulsorincluding a circumferential array of blades; a compressor in fluidcommunication with the propulsor, the compressor including a highpressure compressor section and a low pressure compressor section, andthe low pressure compressor section including a low pressure compressorsection inlet with a low pressure compressor section inlet annulus area;a combustor in fluid communication with the compressor; a shaft assemblyhaving a first portion and a second portion; a two-stage high pressureturbine section coupled to the first portion of the shaft assembly todrive the high pressure compressor section, and a low pressure turbinesection coupled to the second portion of the shaft assembly, each of thehigh pressure turbine section and low pressure turbine section includingblades and vanes, and a low pressure turbine section airfoil countdefined as the numerical count of all of the blades and vanes in the lowpressure turbine section, wherein the low pressure turbine sectionairfoil count is below 1600, wherein the low pressure compressor sectionincludes a greater number of stages than the high pressure turbinesection, and the low pressure compressor section and the low pressureturbine section have an equal number of stages; a speed reductionmechanism coupled to the propulsor and rotatable by the low pressureturbine section through the second portion of the shaft assembly todrive the propulsor; and wherein the low pressure turbine sectionfurther includes a maximum gas path radius, the blades of the propulsorinclude a maximum radius, and a ratio of the maximum gas path radius tothe maximum radius of the blades of the propulsor is equal to or greaterthan 0.35, and is less than 0.55.
 2. The engine as recited in claim 1,wherein a total number of stages of the high pressure compressor sectionis greater than a combined total number of stages of the low pressurecompressor section and the low pressure turbine section.
 3. The engineas recited in claim 2, wherein a hub-to-tip ratio (R_(I):R_(O)) of thelow pressure turbine section is between 0.4 and 0.5 measured at themaximum R_(O) axial location in the low pressure turbine section.
 4. Theengine as recited in claim 3, wherein the speed reduction mechanism isan epicyclic transmission including a sun gear encircled by a pluralityof intermediary gears, a ring gear and a carrier that carries theplurality of intermediate gears, wherein the intermediary gears arepositioned between and enmeshed with the sun gear and the ring gear. 5.The engine as recited in claim 4, wherein the epicyclic transmission hasa speed reduction ratio between 2:1 and 13:1 determined by the ratio ofdiameters within the epicyclic transmission.
 6. The engine as recited inclaim 5, wherein the ratio of the maximum gas path radius to the maximumradius of the array of blades of the propulsor is less than 0.50, andthe hub-to-tip ratio (R_(I):R_(O)) is between 0.42 and 0.48.
 7. Theengine as recited in claim 5, wherein the epicyclic transmission is astar gear system, the sun gear is coupled to a forward end of the secondportion of the shaft assembly, the carrier is mounted to an engine case,and the ring gear is coupled to the propulsor such that the ring gearand the propulsor are rotatable as a unit.
 8. The engine as recited inclaim 5, wherein the epicyclic transmission is a planetary gear system,the sun gear is coupled to a forward end of the second portion of theshaft assembly, the carrier is connected to the propulsor, and the ringgear is fixed to a fixed structure of the engine.
 9. The engine asrecited in claim 8, wherein the low pressure turbine section drives thelow pressure compressor section and an input of the epicyclictransmission.
 10. The engine as recited in claim 9, wherein theepicyclic transmission is axially aft of the low pressure compressorsection inlet annulus area with respect to a longitudinal axis of theengine.
 11. The engine as recited in claim 10, wherein the engine is atwo-spool engine, the propulsor is a fan, and further comprising a fancase encircling the array of blades radially outboard of an engine caseto define a fan duct and a bypass flow path, and the fan duct includinga fan duct annulus area outboard of the low pressure compressor sectioninlet, and the engine having only a single fan stage comprising thearray of blades.
 12. The engine as recited in claim 11, wherein theratio of the maximum gas path radius to the maximum radius of the arrayof blades of the fan is less than 0.50.
 13. The engine as recited inclaim 12, wherein the hub-to-tip ratio (R_(I):R_(O)) is between 0.42 and0.48.
 14. The engine as recited in claim 13, wherein each of the bladesof the low pressure turbine section is an unshrouded blade.
 15. Theengine as recited in claim 13, wherein each of the blades of the lowpressure turbine section includes an airfoil extending from an innerdiameter platform to an outer diameter shroud.
 16. The engine as recitedin claim 13, wherein the low pressure turbine section is a four-stageturbine section.
 17. A gas turbine engine comprising: a propulsorincluding a circumferential array of blades; a compressor in fluidcommunication with the propulsor, the compressor including a nine-stagehigh pressure compressor section and a low pressure compressor section,the low pressure compressor section including a low pressure compressorsection inlet with a low pressure compressor section inlet annulus area,and the low pressure compressor section including four stages; acombustor in fluid communication with the compressor; a shaft assemblyhaving a first portion and a second portion; a two-stage high pressureturbine section coupled to the first portion of the shaft assembly todrive the high pressure compressor section, and a four-stage lowpressure turbine section coupled to the second portion of the shaftassembly, each of the high pressure turbine section and low pressureturbine section including blades and vanes, and a low pressure turbinesection airfoil count defined as the numerical count of all of theblades and vanes in the low pressure turbine section, and wherein thelow pressure turbine section airfoil count is below 1600; and a speedreduction mechanism coupled to the propulsor and rotatable by the lowpressure turbine section through the second portion of the shaftassembly to drive the propulsor; and wherein the low pressure turbinesection further includes a maximum gas path radius, the blades of thepropulsor include a maximum radius, and a ratio of the maximum gas pathradius to the maximum radius of the blades of the propulsor is equal toor greater than 0.35, and is less than 0.55.
 18. The engine as recitedin claim 17, wherein a hub-to-tip ratio (R_(I):R_(O)) of the lowpressure turbine section is between 0.4 and 0.5 measured at the maximumR_(O) axial location in the low pressure turbine section.
 19. The engineas recited in claim 18, wherein the speed reduction mechanism is anepicyclic transmission including a sun gear encircled by a plurality ofintermediary gears, a ring gear and a carrier that carries the pluralityof intermediate gears, wherein the intermediary gears are positionedbetween and enmeshed with the sun gear and the ring gear.
 20. The engineas recited in claim 19, wherein the epicyclic transmission has a speedreduction ratio between 2:1 and 13:1 determined by the ratio ofdiameters within the epicyclic transmission.
 21. The engine as recitedin claim 20, wherein the ratio of the maximum gas path radius to themaximum radius of the array of blades of the propulsor is less than0.50, and the hub-to-tip ratio (R_(I):R_(O)) is between 0.42 and 0.48.22. The engine as recited in claim 20, wherein the epicyclictransmission is a star gear system, the sun gear is coupled to a forwardend of the second portion of the shaft assembly, the carrier is mountedto an engine case, and the ring gear is coupled to the propulsor suchthat the ring gear and the propulsor are rotatable as a unit.
 23. Theengine as recited in claim 20, wherein the epicyclic transmission is aplanetary gear system, the sun gear is coupled to a forward end of thesecond portion of the shaft assembly, the carrier is connected to thepropulsor, and the ring gear is fixed to a fixed structure of theengine.
 24. The engine as recited in claim 23, wherein the low pressureturbine section drives the low pressure compressor section and an inputof the epicyclic transmission.
 25. The engine as recited in claim 24,wherein the epicyclic transmission is axially aft of the low pressurecompressor section inlet annulus area with respect to a longitudinalaxis of the engine.
 26. The engine as recited in claim 25, wherein theengine is a two-spool engine, the propulsor is a fan, and furthercomprising a fan case encircling the array of blades radially outboardof an engine case to define a fan duct and a bypass flow path, and thefan duct including a fan duct annulus area outboard of the low pressurecompressor section inlet, and the engine having only a single fan stagecomprising the array of blades.
 27. The engine as recited in claim 26,wherein the ratio of the maximum gas path radius to the maximum radiusof the array of blades of the fan is less than 0.50.
 28. The engine asrecited in claim 27, wherein the hub-to-tip ratio (R_(I):R_(O)) isbetween 0.42 and 0.48.
 29. The engine as recited in claim 28, whereineach of the blades of the low pressure turbine section is an unshroudedblade.
 30. The engine as recited in claim 28, wherein each of the bladesof the low pressure turbine section includes an airfoil extending froman inner diameter platform to an outer diameter shroud.